l1 Launch System

8/18/2019 l1 Launch System

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Outline

• Launch systems characteristics

• Launch systems selection process

• Spacecraft design envelope &

environments.


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Lesson Objectives• Each student will

• Understand launch system characteristics, sizing and

trade-offs.

• Estimate launch system sizes, staging requirements.

• Be able to select appropriate launch system for a given

mission from available systems.

• Be able to estimate spacecraft requirements driven by

launch vehicle induced environments.

• Determine costs of launch systems.


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Rocket Basics

Thrust

Specific impulse

where K depends on and the engine pressure ratio

)( pe peeV em

cT K Isp

g m

F Isp

me

Ve

Pe

Ae

pFuel in

Oxidizer in

Thrust

pc

Throat


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Rocket Equation

Where is the mass ratio

(Assumes zero losses due to drag and gravity)

t mo M veh M g

veh M

Ispm

veh M

a

V a d t Isp g d md t

d t

M

o

d m

d t

t t o

t f

Rn Isp g f

M o M n Isp g

propellant M o M

o M n Isp g V

R o M

f


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Rocket Equation (Cont.)

M p M f e

V

Isp g

1

M o1eV / Isp g

M p = mass of propellant

Mo = initial mass

Mf = final massV = vehicle velocity changeM

veh

= vehicle mass


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Staging• Near burnout, rocket acceleration is diminished because

payload mass includes entire launch systems structure.

• Staging removes lower stage structural weight

Moi = initial mass of rocket including all upper stages and payload.

Mfi = final mass after stage has burned before separation.

i = stage number

V = Vi

V i g Isp n

M oi M

f i


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Staging (Cont.)

= payload fraction =

or

M payload i = mass of payload plus all upper stages

Structure fraction

Msi = mass structure for stage i

M pi = mass propellant for stage i

payload M o

i

payload

moi

i

i1

n

i

si

si M

oi

si M

pi

M si

M payload i


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Launch Vehicle Forces

W

L

D

V

xz

cp

T

Flight path angle

Angle of attack

ax g T W

sin

D

W cos

L

W sin

az g cos

D

W sin L

W cos


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Example Launch SystemsLaunch

System

Upper

Stage

LEO

(kg)

GTO

(kg)

GEO

(kg)

Polar

(kg)

AtlasIIAS

Centaur 2A

8640 3606 1050 7300

Delta II

7920/25

PAM-D 5089 1840 910 3890

Pegasus

XL

460 345

Shuttle -

IUS

24,400

5900 2360

Taurus Star 37 1400 450 1060

Titan IV -

Centaur 8620 4540

14,110


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Example Orbit

Transfer VehiclesCharacteristics PAM-D IUS Centaur

Length (m) 2.04 5.2 9.0

Diameter (m) 1.25 2.9 4.3

Mass (kg) 2180 14,865 18,800

Thrust (N) 66,440 200,000 147,000

Isp 292.6 292.9 442

Structure mass 180 1255 2100

Propellant mass 2000 9710 16,700

Airborne support

equipment mass

1140 3350 4310


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Launch Sites Criteria• Minimum inclination

• Launch azimuth• Weather Orbital

motion

To sun

Earth rotation


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US Launch Sites and LaunchSystems

• Western range (Vandenberg AFB): – LMMS Titan II, IV-B, Athena

– Boeing Delta II, III, EELV

– OSC Taurus, Minotaur, Pegasus

• Equatorial launch site:

– Boeing SeaLaunch• Alaska Spaceport

– OSC Minotaur


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US Launch Sites and LaunchSystems (continued)

• Eastern Range (Cape Canaveral Air Station,Kennedy Space Center):

– STS

– LMMS Titan IV; Atlas II, IIA, IIAS; EELV – Boeing Delta II, III, EELV

– Orbital Pegasus XL, (Taurus, Minotaur)

– Coleman/TRW/IAI Shavit

• Wallops Island

– Pegasus XL, Minotaur


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Typical Launch VehicleIntegration Tasks

• Mission Orbit Planning – Effect of launch delays, launch window definition

• Launch vehicle and spacecraft performance analyses

– LV performance variations vs mission impacts

• Defining, implementing mission unique requirements – Ground processing, ground testing

– Launch vehicle interfaces - power, command, telemetry, etc.

– Critical s/c commands: self-generated, booster provided, backup

timers?• Flight safety systems - range destruct protocols: installation and test of

range destruct packages

• Developing multi-agency day-of-launch launch ops procedures

– Example: Go/No-Go limits


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Launch Services - Scheduling

• LMMS Atlas Commercial template – @ 36 months, select a 3 month window

– @ 12 months, select a 30 day slot

– @ 6 months, select a launch day

• STS templates:

– 36+ months for a Primary payload – 24 months minimum for secondary payloads


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Payload Integration

• Fairing size and shape

• Maximum accelerations

• Vibration frequencies and magnitudes

• Acoustic frequencies and magnitudes

• Temperature extremes

• Air cleanliness

• Orbital insertion accuracy

• Interfaces to launch site and vehicleGround handling, ground and airborne

transportation, and launch environment may be

more severe than space operating environment


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Fairings

• Protection from aerodynamic loading

• Diameter and length constraint

• Acoustic environment

• Jettison Altitude


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Structural & Electrical Interface

• Physical support adaptors

• Separation/deployment system

• Kick motor/Spin tables

• Electrical interface• Access

– Physical

– Electrical

– Optical

– Radio frequency


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Payload Environments

• Thermal

– Pre-launch

– Ascent fairing radiant – Aero-heating (Free molecular heating)

• Electromagnetic

• Contamination

• Venting

• Acceleration

• Vibration

• Acoustics

• Shock


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Acceleration Load FactorsLift off Max

Airloads

Stage 1

shutdown

Stage 2

shutdown

Vehicle Axial Later al Axial Lateral Axial Later al Axial Later al

Titan 34D/IUS steady

Dynamic

+1.5

+1.5 + 5.0

+2.0

+ 1.0 + 2.5

0 - +4.5

+ 4.0 + 2.0

0 – +2.5

+ 4.0 + 2.0

Atlas II steadydynamic

+1.3+1.5 +1.0

+2.2+ 0.3

+0.4+ 1.2

+5.5+ 0.5 + 0.5

+4.0+ 2.0 0.5

Delta steady

dynamic

+2.4

1.0 2 to 3 +6.0

Shuttle IUS steadyDynamic

+3.2+3.5

+2.5+3.4

+1.1

to 3.2

+0.25 to

–0.59 +3.2 +0.59


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Vibration Environments

• Caused by

– Launch system propulsive dynamic acceleration

– Unsteady aerodynamic effects

– Acoustic pressure from engines

– Amplified mechanical response of vehicle structure• Includes ground and airborne transportation

• Yields structural stiffness requirement on

payload and adaptor/interface.


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Shock Loads

Caused by pyrotechnic devices used to separatefrom launch.

Staging, engine starts and shut down.


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Acoustic Environments

• Caused by

– Reflected sound energy from launch pad structures

and facilities. – Maximum dynamic pressure (max q) aerodynamics.

• Affected by fairing design


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Injection Accuracy• Final stage guidance and propulsion

performance determines injection accuracy. – Apogee, perigee, inclination

– Payload’s Attitude Determination and Control

System must capture and correct linear androtational tip-off rates, and injection errors.


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Payload Integration Procedures

• Mating spacecraft to launch vehicle.

• Spin tests.

• Propellant loading.

• Pre-launch test of all subsystems.


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Payload Processing• Receiving inspection

• Payload & ground support equipment

• Installing hardware (batteries, guidance

systems)

• Pressure checks

• Communication and payload functional test


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Launch System

Cost Estimate

• Determined from supplier.

• Should include integration and check out costs,

launch support systems and launch integration

costs.• Small payloads may ride as a secondary

payload.

• Example launch system costs.


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References – Launch system user handbooks.

– Lockheed Martin, Boeing, Orbital, etc. (or www)

– AIAA Launch Vehicle Summary (in Library)

– International Space Industry Report

– Reducing Space Mission Costs. Wertz and Larson

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